The present invention pertains to the fault detection and indication art and, more particularly, to a system for monitoring fault conditions in any one of several power actuator units used to position the control surface of an aircraft.
Modern aircraft commonly employ multiple hydraulic power actuators to position aricraft control surfaces, such as the elevators or rudder. FIG. 1 is a cutaway, perspective view of an aircraft stabilizer portion 10. Shown is a portion of a conventional elevator 12, which is driven into position by three parallel power servo control units 20, 30 and 40. The input command to each control unit is provided by a pilot's input rod 50. Input rod 50 translates, in the directions indicated, in response to a flight command from the flight deck.
The translation motion of the pilot's input rod 50 causes a rotation of bellcranks 21, 31 and 41. Rotation of the bellcranks 21, 31 and 41 results in a translation of the adjustment links 22, 32 and 42. This creates a rotation of summing levers 23, 33, 43 and a corresponding translation of the first power servo input links 24, 34 and 44. Such translation produces a rotation of the second power servo input links 25, 35 and 45. The second links 25, 35 and 45 connect to the control inputs of conventional hydraulic actuators 26, 36, 46. Pistons within the actuators 26, 36, 46 respond to changes in hydraulic fluid pressure and fluid flow caused by control inputs to deflect corresponding piston rods 27, 37, 47. The piston rods connect to a shaft 52 which, via brackets, such as bracket 54, couple to the elevator 12. The summing levers 23, 33, 43 are pivotally conneted to the piston rods 27, 37, 47, thereby providing a mechanical servo system to assure that each piston rod 27, 37, 47 is positioned proportionally to the translation of the pilot's input rod 50.
Thus, all power servo control units 20, 30, 40 receive essentially identical input commands via the pilot's input rod 50. In the present example, the units 20, 30, 40 are sized comparably such that each shares equally in the load on the elevator 12.
In addition, if any one of the units 20, 30, 40 fails the remaining functioning units are capable of overpowering the failed unit, thereby maintaining position control of the elevator surface 12. Such failures are typically passive in nature and would normally go unnoticed by the flight crew until such time as a considerable amount of control power was required to position elevator 12, or until an additional failure in one of the power servo control units occurred. The condition of excess power required to control the elevator 12 might result in reduced elevator control, thereby reducing flight control of the aircraft. In the event of a second power servo control unit failure, a similar flight problem could occur. It is desirable, therefore, to detect first failures promptly, such that they may be corrected prior to the development of a more serious condition.
The power servo control units are subject to three failure modes of primary concern:
jamming of the control valve;
a disconnect between the control valve and the control surface; and,
a disconnect or shearout of the control linkage.
The prior art has developed numerous systems to detect a failure in a power servo control unit. In one such system, a valve jam causes a detented bundee in the control linkage to move, thereby actuating a switch and annunciating the condition. In a second approach, a valve jam causes a detented valve within the primary control valve to port fluid to a pressure switch, thereby annunciating the failure. In a third prior art approach, upon valve jam, the detented bungee in the control linkage moves, thereby changing the normal control feel forces in the system, allowing pilot detection.
In all of these prior art systems, only valve jams are detected, all other failures of concern being detected only upon inspection.
An additional problem with such prior art failure detection systems is that they do not provide a means to detect a misrigging, or other power servo control unit overloading condition, which, although not representing a failure of the servo control unit, nonetheless results in reduced useful life of the servo control unit system and structure. For example, referring to FIG. 1, during rigging of the power servo control units 20, 30, 40, adjustments are made on links 22, 32, 42 to minimize mistracking among the power servo control units. Such mistracking leads to a "force fight" among the units. The failure to properly set these adjustment links, or misrigging caused by other conditions such as damage to one of the linkage members, may result in a substantial mistracking of the servo control units. This, it has been shown, results in the aforementioned "force fight" with a reduced expected lifetime of the servo control units and structure.
There is a need in this art, therefore, for a fault monitoring system which is capable of detecting and indicating any type of excessive load producing fault in a power servo control unit. In addition, there is a need in this art for such a monitoring and indicating system which is capable of detecting misrigging or other mistracking conditions which may lead to reduced component life.